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NACA M8 AIRFOIL (m8-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA M8 AIRFOIL (m8-il)
Reynolds number: 100,000
Max Cl/Cd: 40.38 at α=7.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m8-il-100000.txt
Download as CSV file: xf-m8-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M8 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.1491   0.11029   0.10608  -0.0236   0.8639   0.0787
  -9.000  -0.1541   0.11107   0.10684  -0.0262   0.8474   0.0801
  -8.750  -0.1667   0.11283   0.10860  -0.0291   0.8323   0.0805
  -8.500  -0.1343   0.10346   0.09915  -0.0251   0.8192   0.0823
  -8.250  -0.1235   0.10071   0.09635  -0.0244   0.8057   0.0848
  -8.000  -0.1172   0.09866   0.09423  -0.0243   0.7945   0.0875
  -7.750  -0.1124   0.09698   0.09255  -0.0254   0.7830   0.0903
  -7.500  -0.1216   0.09805   0.09364  -0.0283   0.7733   0.0921
  -7.250  -0.1216   0.09741   0.09300  -0.0325   0.7642   0.0928
  -7.000  -0.0996   0.09081   0.08634  -0.0280   0.7566   0.0945
  -6.750  -0.0863   0.08803   0.08355  -0.0277   0.7474   0.0971
  -6.500  -0.0780   0.08614   0.08160  -0.0282   0.7407   0.1009
  -6.250  -0.0712   0.08827   0.08371  -0.0396   0.7314   0.1053
  -6.000  -0.0590   0.08233   0.07776  -0.0342   0.7250   0.1068
  -5.750  -0.0441   0.07946   0.07490  -0.0335   0.7169   0.1095
  -5.500  -0.0301   0.07738   0.07278  -0.0347   0.7093   0.1138
  -5.250  -0.0122   0.07636   0.07166  -0.0419   0.7025   0.1186
  -5.000   0.0011   0.07288   0.06822  -0.0393   0.6940   0.1211
  -4.750   0.0220   0.07185   0.06704  -0.0433   0.6877   0.1293
  -4.500   0.0378   0.06873   0.06398  -0.0433   0.6789   0.1320
  -4.250   0.0560   0.06666   0.06184  -0.0437   0.6721   0.1376
  -3.750   0.0988   0.06252   0.05760  -0.0475   0.6568   0.1477
  -3.500   0.1232   0.06049   0.05541  -0.0500   0.6514   0.1531
  -3.000   0.1724   0.05657   0.05137  -0.0537   0.6356   0.1635
  -2.750   0.1953   0.05464   0.04940  -0.0544   0.6280   0.1679
  -2.500   0.2325   0.05345   0.04797  -0.0592   0.6200   0.1726
  -2.250   0.2499   0.05062   0.04506  -0.0579   0.6151   0.1748
  -2.000   0.2761   0.04893   0.04337  -0.0593   0.6051   0.1783
  -1.750   0.3058   0.04705   0.04127  -0.0607   0.5991   0.1802
  -1.500   0.3346   0.04531   0.03941  -0.0624   0.5911   0.1798
  -1.250   0.3732   0.04125   0.03493  -0.0665   0.5846   0.1494
  -1.000   0.4134   0.03489   0.02774  -0.0703   0.5807   0.1299
  -0.750   0.4362   0.03540   0.02841  -0.0695   0.5722   0.1340
  -0.500   0.4670   0.03340   0.02602  -0.0701   0.5661   0.1402
  -0.250   0.4929   0.03298   0.02547  -0.0690   0.5616   0.1456
   0.000   0.5195   0.03283   0.02529  -0.0691   0.5520   0.1550
   0.250   0.5473   0.03216   0.02439  -0.0682   0.5460   0.1664
   0.500   0.5754   0.03189   0.02390  -0.0679   0.5383   0.1837
   0.750   0.5997   0.03166   0.02365  -0.0671   0.5306   0.2009
   1.000   0.6270   0.03092   0.02259  -0.0662   0.5257   0.2306
   1.250   0.6496   0.03066   0.02239  -0.0659   0.5165   0.2589
   1.500   0.6748   0.02970   0.02126  -0.0651   0.5106   0.2913
   1.750   0.7017   0.02907   0.02036  -0.0644   0.5059   0.3224
   2.000   0.7241   0.02889   0.02027  -0.0642   0.4978   0.3436
   2.250   0.7510   0.02844   0.01965  -0.0636   0.4923   0.3603
   2.500   0.7783   0.02851   0.01952  -0.0631   0.4870   0.3707
   2.750   0.8015   0.02857   0.01968  -0.0629   0.4791   0.3826
   3.000   0.8291   0.02819   0.01917  -0.0624   0.4739   0.3950
   3.250   0.8555   0.02838   0.01928  -0.0621   0.4680   0.3990
   3.500   0.8806   0.02865   0.01958  -0.0618   0.4602   0.3963
   3.750   0.9106   0.02836   0.01908  -0.0614   0.4551   0.3913
   4.000   0.9344   0.02901   0.01974  -0.0610   0.4470   0.3866
   4.250   0.9620   0.02883   0.01950  -0.0606   0.4402   0.3841
   4.500   0.9906   0.02863   0.01916  -0.0602   0.4347   0.3822
   4.750   1.0134   0.02912   0.01976  -0.0598   0.4255   0.3800
   5.000   1.0451   0.02876   0.01919  -0.0596   0.4200   0.3764
   5.250   1.0673   0.02942   0.01996  -0.0592   0.4103   0.3731
   5.500   1.0977   0.02916   0.01950  -0.0588   0.4038   0.3700
   5.750   1.1214   0.02965   0.02008  -0.0586   0.3941   0.3681
   6.000   1.1523   0.02936   0.01963  -0.0585   0.3870   0.3658
   6.250   1.1743   0.03002   0.02038  -0.0580   0.3775   0.3639
   6.500   1.2030   0.03000   0.02022  -0.0578   0.3706   0.3618
   6.750   1.2241   0.03082   0.02113  -0.0572   0.3625   0.3601
   7.000   1.2485   0.03124   0.02151  -0.0567   0.3554   0.3585
   7.250   1.2761   0.03160   0.02172  -0.0563   0.3500   0.3572
   7.500   1.2902   0.03292   0.02328  -0.0553   0.3423   0.3562
   7.750   1.3160   0.03325   0.02355  -0.0549   0.3371   0.3554
   8.000   1.3385   0.03404   0.02432  -0.0543   0.3321   0.3556
   8.250   1.3491   0.03553   0.02606  -0.0530   0.3259   0.3560
   8.500   1.3707   0.03617   0.02671  -0.0523   0.3213   0.3569
   8.750   1.3997   0.03654   0.02697  -0.0521   0.3179   0.3580
   9.000   1.4044   0.03869   0.02940  -0.0507   0.3136   0.3584
   9.250   1.4069   0.04089   0.03186  -0.0491   0.3093   0.3586
   9.500   1.4181   0.04243   0.03351  -0.0480   0.3058   0.3588
   9.750   1.4392   0.04333   0.03443  -0.0474   0.3030   0.3594
  10.000   1.4698   0.04387   0.03490  -0.0474   0.3006   0.3606
  10.250   1.4249   0.04938   0.04085  -0.0438   0.2975   0.3597
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